Product made of an AlCuMg alloy for aircraft structural elements

ABSTRACT

Rolled, extruded or forged product made of an AlCuMg alloy processed by solution heat treatment, quenching and cold stretching, to be used in the manufacture of aircraft structural elements, with the following composition (% by weight):  
     Fe&lt;0.15 Si&lt;0.15 Cu:3.8-4.4 Mg:1-1.5 Mn:0.5-0.8 Zr:0.08-0.15  
     other elements: &lt;0.05 each and &lt;0.15 total with a ratio R m (L)/R 0.2 (L)&gt;1.25.  
     The invention is particularly applicable to the manufacture of lower wings, and has a set of properties (toughness, crack propagation rate, fatigue strength, residual stress level), that are better than alloy 2024.

TECHNICAL FIELD

[0001] This invention relates to rolled, extruded or forged productsmade of a quenched and stretched AlCuMg alloy designed for themanufacture of aircraft structural elements, particularly skin panelsand lower wing stringers, and with an improved compromise betweenmechanical strength, formability, toughness, tolerance to damage andresidual stress properties than is available with products according toprior art used for the same application. Designations of alloys andmetallurgical tempers are in accordance with the terminology used by theAluminum Association, and repeated in European standards EN 515 and EN573.

TEMPER OF PRIOR ART

[0002] Wings for high capacity commercial aircraft comprise an upperpart (upper wing) composed of a skin made of thick 7150 alloy plates intempers T651, or 7055 alloy plates in temper T7751 or 7449 alloy platesin temper T7951, and stringers made from profiles of the same alloy, anda lower part (lower wing) composed of a skin made of thick plates madeof a 2024 alloy in temper T351 or a 2324 alloy in temper T39, andstringers made from profiles of the same alloy. The two parts areassembled by spars and ribs.

[0003] The chemical composition of 2024 alloy according to theterminology used by the Aluminum Association and standard EN 573-3 is asfollows (% by weight):

[0004] Si<0.5 Fe<0.5 Cu:3.8-4.9 Mg:1.2-1.8 Mn:0.3-0.9 Cr<0.10 Zn<0.25Ti<0.15

[0005] Different variants have been developed and registered with theAluminum Association as 2224, 2324 and 2424, particularly with lowercontents of silicon and iron. Alloy 2324 in temper T39 was described inBoeing patent EP 0038605 (=U.S. Pat. No. 4,294,625), in which theimprovement to the yield stress is obtained by work hardening by a coldrolling pass after quenching. This work hardening tends to reduce thetoughness, and the contents of Fe, Si, Cu and Mg are reduced in order tocompensate the drop in toughness. Boeing also developed alloy 2034 withcomposition:

[0006] Si<0.10 Fe<0.12 Cu:4.2-4.8 Mg:1.3-1.9 Mn:0.8-1.3 Cr<0.05 Zn<0.20Ti<0.15 Zr:0.08-0.15

[0007] This alloy is described in patent EP 0031605 (=U.S. Pat. No.4,336,075). Compared with alloy 2024 in temper T351, it has a betterspecific yield stress due to an increase in the manganese content andthe addition of another anti-recrystallizing agent (Zr), and also hasimproved toughness and fatigue strength.

[0008] Alcoa patent EP 0473122 (=U.S. Pat. No. 5,213,639) describes analloy, recorded by the Aluminum Association as 2524, with composition:

[0009] Si<0.10 Fe<0.12 Cu:3.8-4.5 Mg:1.2-1.8 Mn:0.3-0.9 that can alsocontain an other anti-recrystallizing agent (Zr, V, Hf, Cr, Ag or Sc).This alloy is intended specifically for thin plates for fuselages andits toughness and resistance to crack propagation are improved comparedwith 2024.

[0010] Patent application EP 0731185 made by the applicant relates to analloy, subsequently registered as 2024A with composition:

[0011] Si<0.25 Fe<0.25 Cu:3.5-5 Mg:1-2 Mn<0.55 satisfying the relation:0<Mn-2 Fe<0.2

[0012] The thick plates made of this alloy are tougher and residualstresses are lower, without any loss of other properties.

[0013] Alcoa patents U.S. Pat. No. 5,863,359 and U.S. Pat. No. 5,865,914relate to an aircraft wing with a lower wing made of an alloy withcomposition:

[0014] Cu:3.6-4 Mg:1-1.6 (preferably 1.15-1.5) Mn:0.3-0.7 preferably0.5-0.6), Zr:0.05-0.25 and

[0015] preferably Fe<0.07 and Si<0.05 with the following properties:

[0016] R_(0.2)(LT)>60 ksi (414 MPa) and K_(ie) (L-T)>38 ksi{squareroot}inch (42 MPa{square root}m),

[0017] and a process for manufacturing a lower wing element withR_(0.2)(LT)>60 ksi comprising casting of an alloy with the previouscomposition, homogenization between 471 and 482° C., hot transformationat a temperature>399° C.; solution heat treatment above 488° C.,quenching cold work hardening preferably by more than 9% and stretchingby at least 1%.

PROBLEM THAT ARISES

[0018] One essential constraint when constructing new high capacitycommercial aircraft is to limit the weight, such that manufacturerspecifications impose higher typical stresses for wing panels, whichleads to higher minimum values for static mechanical properties andhigher damage tolerance for the aluminum alloy products used. The use ofwork hardened products in temper T39, as recommended in patents U.S.Pat. No. 5,863,359 and U.S. Pat. No. 5,865,914, does give higher yieldstresses R_(0.2), but it also has a number of disadvantages for otherworking properties that are important in this application. One result isa very small plastic range, in other words the difference between theultimate stress R_(m) and the yield stress R_(0.2), which results inlower cold formability and less resistance to crack propagation under aload with a variable amplitude. The reduction in the rate of crackpropagation after a partial overload is less important if the plasticrange is small.

[0019] Furthermore, large parts must be machined without distortion inthicker plates, which requires better control of residual stresses.However, temper T39 is not particularly advantageous from this point ofview.

[0020] Therefore the purpose of this invention is to provide AlCuMgalloy products in the work hardened and cold deformed temper, for use inmanufacturing aircraft lower wings, and providing a better compromise ofall working properties (mechanical strength, rate of crack propagation,toughness, resistance to fatigue and residual stresses) than is possiblewith similar products according to prior art.

PURPOSE OF THE INVENTION

[0021] The purpose of the invention is a rolled, extruded or forgedproduct made of an AlCuMg alloy processed by solution heat treatment,quenching and cold stretching, to be used in the manufacture of aircraftstructural elements, with the following composition (% by weight):

[0022] Fe<0.15 Si<0.15 Cu:3.8-4.4(preferably: 4.0-4.3) Mg:1.0-1.5Mn:0.5-0.8 Zr:0.08-0.15

[0023] other elements: <0.05 each and <0.15 total with a ratioR_(m)(L)/R_(0.2)(L) of the ultimate tensile strength in the L directionto the yield stress in the L direction exceeding 1.25 (and preferablyexceeding 1.30).

[0024] Another purpose is a rolled product (plate) with the samecomposition and between 6 and 60 mm thick and with at least thefollowing groups of properties in the quenched and stretched temper:

[0025] a) Ultimate tensile strength R_(m(L))>475 MPa and yield stressR_(0.2(L))>370 MPa

[0026] b) Plastic range R_(m)-R_(0.2) in the L and TL directions>100MPa.

[0027] c) Critical intensity factor (L-T direction) K_(c)>170 MPa{squareroot}m and K_(co)>120 MPa{square root}m (measured according to ASTMstandard E 561 on notched test pieces sampled at a quarter thicknesswith parameters B=5 mm, W=500 and 2_(B0)=165 mm)

[0028] d) Crack propagation rate (L-T direction) da/dn, measuredaccording to ASTM standard E 647 on notched test pieces sampled at aquarter thickness with parameters W=200 mm and B=5 mm)

[0029] <10⁻⁴ mm/cycle for ΔK=10 MPa{square root}m

[0030] <2.5 10⁻⁴ mm/cycle for ΔK=15 MPa{square root}m

[0031] and <5 10⁻⁴ mm/cycle for ΔK=20 MPa{square root}m

[0032] This plate also has residual stresses such that the deflection fmeasured in the L and TL directions after machining a bar supported ontwo supports separated by a length l to its mid-thickness, is such that:

[0033] f<(0.14 l²)/e, where f is measured in microns, e is the thicknessof the plate and l is the length measured in mm.

[0034] Another purpose of the invention is a process for manufacturing arolled, extruded or forged product comprising the following steps:

[0035] cast a plate or billet with the indicated composition,

[0036] homogenize this plate or billet between 450 and 500° C.,

[0037] hot transformation, and possibly cold transformation, until therequired product is obtained,

[0038] solution heat treatment at a temperature of between 480 and 505°C.,

[0039] quench in cold water,

[0040] cold stretching to at least 1.5% permanent deformation,

[0041] natural aging under ambient conditions.

DESCRIPTION OF THE INVENTION

[0042] The chemical composition of the product is different from thechemical composition of the usually used 2024 in that the iron andsilicon contents are lower, the manganese content is higher, andzirconium is added. Compared with 2034, the magnesium content is lowerand the copper content is slightly lower. Compared with the compositionof the alloys described in patents U.S. Pat. No. 5,863,359 and U.S. Pat.No. 5,865,914, the copper content is higher, compensating the lower workhardening after quenching, for the mechanical strength. Surprisingly,this narrow composition range (particularly for manganese) combined withmodifications to the manufacturing procedure, can give a significantimprovement in the compromise between the mechanical strength,elongation and damage tolerance under operating conditions for a highcapacity civil aircraft, compared with prior art. Furthermore, and quiteunexpectedly, low residual stresses are observed particularly for thickproducts, so that large parts can be machined without distortion.

[0043] The manufacturing process consists of casting the plates in thecase in which the product to be made is a rolled plate, or the billetsin the case of an extruded profile or forged part. The plate or thebillet is scalped and then homogenized at between 450 and 500° C. Thehot transformation is then made by rolling, extrusion or forging. Thistransformation is preferably made at a temperature higher thantemperatures normally used, the output temperature being greater than420° C. and preferably greater than 440° C. so that the treated producthas a slightly recrystallized structure, with a recrystallization rateof less than 20%, and preferably less than 10%, at a quarter thickness.The rolled, extruded or forged semi-product is then put into solutionheat treatment at between 480 and 505° C., such that solution heattreatment is as complete as possible, in other words that the maximumnumber of potentially soluble phases, and particularly Al₂Cu and Al₂CuMgprecipitates, are actually in solid solution. The quality of thesolution heat treatment may be evaluated by differential enthalpicanalysis (AED) by measuring the specific energy using the area of thepeak on the thermogram. This specific energy must preferably be lessthan 2 J/g.

[0044] Quenching is then done with cold water, followed by controlledstretching to give a permanent elongation of not less than 1.5%.Finally, the product is aged naturally at ambient temperature.

[0045] Products according to the invention have significantly improvedstatic mechanical properties compared with alloy 2024-T351, currentlyused for aircraft lower wings, and only slightly lower than theproperties of 2034-T351. The high plastic range and elongation of thematerial give excellent cold formability. The toughness, measured bycritical stress intensity factors in plane stress K_(c) and K_(co) isgreater than the toughness of 2024 and 2034 by more than 10%, and thecrack propagation rate da/dn is significantly better than these twoalloys, particularly for high values of ΔK, and for loads with variableamplitude. Fatigue lives measured on notched samples taken atmid-thickness in the L direction are also more than 20% better than with2024 and 2034. Finally, the magnitude of residual stresses measured bythe deflection f after machining a bar supported on two supportsseparated by distance l to half its thickness, is fairly low, althoughthe opposite might have been expected with a fibrous structure. Thisdeflection, measured in microns, is always less than the quotient (0.14l²)/e, where the length l and the thickness e of the plate are expressedin mm.

[0046] All these properties mean that products according to theinvention are particularly suitable for manufacturing aircraftstructural elements, particularly lower wings, but also profiles for awing spar box, assembled spar booms and rib flanges and fuselage skinsand stringers.

EXAMPLES

[0047] Three 1450 mm wide and 446 mm thick plates were cast, made of2024, 2034 alloys and the alloy according to the invention,respectively. The chemical compositions (% by weight) of the alloys areas given in Table 1: TABLE 1 Alloy Si Fe Cu Mg Mn Zr 2024 0.12 0.20 4.061.36 0.54 0.002 2034 0.05 0.07 4.30 1.34 0.98 0.104 Invention 0.06 0.084.14 1.26 0.65 0.102

[0048] The plates were scalped, and then homogenized under the followingconditions:

[0049] For the 2024, 2 h at 495° C. then 5 h at 460° C.

[0050] For the 2034, 5 h at 497° C.

[0051] For the alloy according to the invention, the temperature wasincreased in 12 h and kept for 6 h at 483° C.

[0052] A part of the plates was then hot rolled to a thickness of 40 mmby successive passes of the order of 20 mm. Another part of the plateswas hot rolled to 15 mm. For the alloy according to the invention, thehot rolling entry temperature was 467° C., the exit temperature at 40 mmwas equal to 465° C. and at 15 mm was 444° C.

[0053] Plates were put into solution heat treatment under the followingconditions:

[0054] 3 h and 6 h at 497° C. for 2024 plates with thicknesses equal to15 and 40 mm respectively,

[0055] 2 h and 5 h at 499° C. for 2034 plates with thicknesses equal to15 and 40 mm,

[0056] 9 h at 497° C. for plates according to the invention.

[0057] After quenching in cold water, all plates were subjected tocontrolled stretching to give a permanent elongation of 2%.

[0058] The static mechanical characteristics in the L and TL directionswere then measured, to determine the ultimate stress R_(m) (in MPa), theconventional yield stress at 0.2% R_(0.2) (in MPa) and elongation atfailure A (in %). The results are given in table 2: TABLE 2 AlloyThickness Direction R_(m) R_(0.2) A 2024 40 L 468 362 20.0 2024 40 TL469 330 17.4 2024 15 L 462 360 21.2 2024 15 TL 467 325 17.6 2034 40 L534 416 11.2 2034 40 TL 529 393 12.0 2034 15 L 548 431 13.8 2034 15 TL531 395 14.6 Invention 40 L 510 384 15.4 Invention 40 TL 475 336 18.9Invention 15 L 501 390 16.7 Invention 15 TL 491 351 19.1

[0059] The toughness was also measured by critical intensity factors inplane stress K_(c) and Ke₀ (in MPa{square root}m) in the L-T directionaccording to ASTM standard E 561, on CCT test pieces sampled at quarterthickness, with width W=500 mm and thickness B=5 mm, and a central notchmachined by electroerosion 2a₀=165 mm, enlarged to 170 mm by a fatiguetest. Table 3 contains the results: TABLE 3 Alloy Thickness K_(c) K_(c0)2024 40 143.4 105.2 2034 40 128.8 97.8 Invention 40 179.7 122 2034 15136.4 103.7 Invention 15 173.6 124.3

[0060] The fatigue crack propagation rate da/dn was also measured in theL-T direction (in mm/cycle) for different values of the ΔK (inMPa{square root}m ) according to ASTM standard E 647. This was doneusing two CCT samples with width W=200 mm and thickness B=5 mm, sampledat a quarter-plate thickness in the L-T direction. The length of thecentral notch machined by electroerosion is 30 mm, and this notch isenlarged by the fatigue test to 40 mm. The crack rate measurement testis carried out on an MTS machine with a load at R=0.05 and a stress of40 MPa, calculated to give a value of ΔK equal to 10 MPa{square root}mfor the initial notch length of 40 mm (results in table 4). TABLE 4Alloy Thickness ΔK = 10 ΔK = 12 ΔK = 15 ΔK = 20 ΔK = 25 2024 40 9 10⁻⁵1.5 10⁻⁴ 3.0 10⁻⁴ 6 10⁻⁴ 9 10⁻⁴ 2034 40 8 10⁻⁵ 1.5 10⁻⁴ 3 10⁻⁴ 5.7 10⁻⁴1.7 10⁻³ Invention 40 5.5 10⁻⁵ 1.7 10⁻⁴ 2.0 10⁻⁴ 4.0 10⁻⁴ 7.8 10⁻⁴ 203415 8 10⁻⁵ 1.5 10⁻⁴ 3 10⁻⁴ 5.2 10⁻⁴ 2.1 10⁻³ Invention 15 4.9 10⁻⁴ 6.010⁻⁴ 1.3 10⁻⁴ 2.5 10⁻⁴ 5.4 10⁻⁴

[0061] Fatigue tests according to the Airbus specification AITM 1-0011were carried out on 7.94 mm thick perforated test pieces 230 mm long, 50mm wide, sampled at mid-thickness in the plate in the L direction. Thehole diameter is 7.94 mm. An average stress of 80 MPa on the solid testpiece was applied with four alternating stress levels: 85 MPa, 55 MPa,45 MPa and 35 MPa for 40 mm plates, and with stresses of 110, 85, 55 and45 MPa for 15 mm plates, with 2 test pieces per level. The average lifevalues (as a number of cycles) are given in table 5. It is found thatthe fatigue life is more than 20% better than with alloy 2024, with anotch factor K_(t)=2.5. TABLE 5 Thickness 80 ± 85 80 ± 55 80 ± 45 80 ±35 Alloy mm MPa MPa MPa MPa 2024 40 36044 159721 2034 40 30640 125565340126  839340 Invention 40 42933 219753 392680 1018240 2034 15 41040204038 352957 Invention 15 45841 241932 429895

[0062] Finally, the deflections f in the L and TL direction weremeasured, together with the recrystallization rate (in %) at thesurface, at a quarter thickness and at half-thickness, determined byimage analysis after chemical etching of the sample.

[0063] The deflection f is measured as follows. Two bars are taken fromthe plate with thickness e, one called the L direction bar with length bin the direction of the length of the plate (L direction), 25 mm wide inthe direction of the width of the plate (TL direction) and withthickness e equal to the full thickness of the plate (TC direction), theother bar being called the TL direction bar with dimensions of 25 mm inthe L direction, b in the TL direction and e in the TC direction.

[0064] Each bar is machined down to half-thickness and the deflection atmid-length of the bar is measured. This deflection is representative ofthe internal stresses in the plate and its ability to not deform duringmachining. The distance l between supports was 180 mm and the length bof the bars was 200 mm. Machining was done mechanically andprogressively with passes of about 2 mm. The deflection at mid-lengthwas measured using a dial gauge with a resolution of one micron. Theresults of the deflections and recrystallization rates are shown intable 6. TABLE 6 Work Work Work hard. hard. hard. f_(L) f_(TL) ratioratio ratio Alloy Thickness (μm) (μm) (surf.) % (¼ t) % (½ t) % 2024 40210 120 79 58  30  2034 40 147 129 12 0 0 Invention 40  86  75 46 5 2

1. Rolled, extruded or forged product made of an AlCuMg alloy processedby solution heat treatment, quenching and cold stretching, to be used inthe manufacture of aircraft structural elements, with the followingcomposition (% by weight): Fe<0.15 Si<0.15 Cu:3.8-4.4 (preferably:4.0-4.3) Mg:1-1.5 Mn:0.5-0.8 Zr:0.08-0.15 other elements: <0.05 each and<0.15 total with a ratio R_(m)(L)/R_(0.2)(L)>1.25 (andpreferably >1.30).
 2. Product according to claim 1, characterized inthat Fe+Si<0.15%
 3. Rolled product between 6 and 60 mm thick accordingto one of claims 1 or 2, with an ultimate tensile strength R_(m(L)) inthe quenched and stretched temper>475 MPa and yield stressR_(0.2)(L)>370 MPa.
 4. Rolled product between 6 and 60 mm thickaccording to any one of claims 1 to 3, with a plastic range between theultimate tensile strength R_(m) and the yield stress R_(0.2) in the Land TL directions in the quenched and stretched temper>100 MPa. 5.Rolled product between 6 and 60 mm thick according to any one of claims1 to 4, for which the critical intensity factor (L-T direction) K_(c) inthe quenched and stretched temper>170 MPa{square root}m and K_(co)>120MPa{square root}m measured according to ASTM standard E 561 on notchedtest pieces sampled at a quarter thickness with parameters W=500 mm, B=5mm and 2_(B0)=165 mm).
 6. Rolled product according to any one of claims1 to 5, characterized in that it has a critical intensity factor (L-Tdirection) K_(c) or K_(c0) at least 10% higher than alloy 2024 under thesame conditions.
 7. Rolled product according to any one of claims 1 to6, characterized in that the crack propagation rate (L-T direction)da/dn in the quenched and stretched temper, measured according to ASTMstandard E 647 on notched test pieces sampled at a quarter thicknesswith parameters W=200 mm and B=5 mm) is as follows: <10⁻⁴ mm/cycle forΔK=10 MPa{square root}m <2.5 10⁻⁴ mm/cycle for ΔK=15 MPa{square root}mand <5 10⁻⁴ mm/cycle for ΔK=20 MPa{square root}m
 8. Rolled productaccording to any one of claims 1 to 7, characterized in that deflectionf is measured in the L and TL directions after machining a bar supportedon two supports separated by a length l to its mid-thickness<(0.14l²)/e, where f is measured in microns, e is the thickness of the plateand l is the length measured in mm.
 9. Rolled product according to anyone of claims 3 to 8, characterized in that the average fatigue lifemeasured on a notched sample taken at mid-thickness in the L directionis more than 20% better than with the 2024 alloy.
 11. Process accordingto any one of claims 1 to 9 for manufacturing a product comprising thefollowing steps: cast a plate with the indicated composition, homogenizethis plate between 450 and 500° C., hot transformation, and possiblycold transformation by rolling, extrusion or forging, until the requiredproduct is obtained, solution heat treatment at a temperature of between480 and 505° C., quench in cold water, cold stretching to at least 1.5%permanent deformation, natural aging under ambient conditions. 12.Process according to claim 10, characterized in that the hottransformation takes place with an exit temperature>420° C. andpreferably >440° C.
 13. Use of plates according to one of claims 3 to 9for manufacturing the skin of an aircraft lower wing.
 14. Use ofprofiles according to one of claims 1 or 2, for manufacturing aircraftlower wings or fuselage stringers.